Most conventional communication spacecraft comprise two solar array panels which extend from opposite surfaces of the spacecraft. These solar array panels are employed to power spacecraft loads (e.g., communications hardware) during time periods when the spacecraft is exposed to rays of the sun. The solar arrays also generate energy that is subsequently stored within batteries located on the spacecraft. These batteries are used to power the spacecraft loads during time periods ("dark periods") when the spacecraft is not exposed to solar energy.
Conventional communication spacecraft also include power electronics circuits for regulating output voltages of the solar array panels and batteries to predetermined voltage levels. After regulating the output voltages of these devices, the power electronics circuits distribute power to designated spacecraft loads.
FIG. 1a shows an example of a single power bus system 1 for a spacecraft. The system 1 includes a solar array 2, a battery 4, and a power electronics block or circuit 6. The power electronics block 6 has an output that is coupled to a spacecraft electrical load (L) via a bus 8. The solar array 2 powers the load (L) during periods when the solar array 2 is exposed to rays of the sun, and the battery 4 is used to power the load (L) during dark periods. The system 1 suffers from a disadvantage that it does not provide a high degree of performance reliability (i.e., fault tolerance). For example, in the event that one or more of the components of the system 1 become short circuited, the spacecraft load (L) can experience a total power loss.
An example of a conventional dual power bus system 3 for a space craft is shown in FIG. 1b. The system 3 includes two single power bus systems that are labelled "9a" and "9b". Each of the systems 9a and 9b is similar to the system 1 of FIG. 1a. An output of the system 9a is connected through bus 8 to an input 11a of a switch or relay 11, and an output of system 9b is connected through another bus 8 to an input 11b of the switch 11. Outputs 11c and 11d of the switch 11 are connected to electrical loads (L.sub.A) and (L.sub.B), respectively, via a set of primary buses 10 and 12. The switch 11 is controllable by a controller (not shown) in a known manner for coupling power output by the individual systems 9a and 9b to selected ones of the loads (L.sub.A) and (L.sub.B) of the spacecraft.
The dual power bus system 3 can provide a greater level of performance reliability for powering spacecraft loads than can be provided by the single power bus system 1. This can be understood in view of the following exemplary situation. In this exemplary situation, it is assumed that while the systems 9a and 9b of the system 3 are being used to power the loads (L.sub.A) and (L.sub.B) , respectively, the bus 10 becomes short circuited and, as a result, no power is provided to the load (L.sub.A). It is also assumed that the system 9b remains unaffected by the short circuit that occurs to the bus 10. As can be appreciated, although power is no longer provided to the load (L.sub.A) due to the short circuit, power is still provided to the load (L.sub.B) from system 9b, and thus there is no total loss of the spacecraft loads. Therefore, assuming that at least a portion of the load (L.sub.B) includes hardware for communicating with a ground station, communications can still be effected between the spacecraft and the ground station.
Although the system 3 can provide a greater level of performance reliability than the system 1 of FIG. 1a, the level of performance reliability that can be provided by the system 3 is nevertheless limited. By example, for a case in which both of the primary buses 10 and 12 become short circuited, both of the loads (L.sub.A) and (L.sub.B) may experience a loss of power. It can therefore be appreciated that, although the prior art power systems 1 and 3 may be suitable for their intended applications, they may not be suitable in other applications which require a greater degree of system performance reliability than can be provided by the systems 1 and 3. Therefore, it would be desirable to provide a spacecraft power system that can provide a greater level of performance reliability than is provided by these conventional spacecraft power systems.
The prior art power systems 1 and 3 also suffer from another shortcoming. Regulated output voltage levels provided by typical, commercially-available power electronics circuits such as, for example, the blocks 6 of the systems 1 and 3, are generally about 28 volts .+-.2 volts or 100 volts .+-.2 volts. Unfortunately, in some cases in which these power electronics circuits are employed in the systems 1 and 3, such power levels are insufficient for supporting high spacecraft load demands. As a result, a need can arise for designing and manufacturing power electronics circuits that can provide higher regulated power levels for these systems 1 and 3. Unfortunately, however, it can be expensive and time consuming to design and manufacture such higher power regulation circuits. Therefore, it would be desirable to provide a spacecraft power system that can power spacecraft loads having high power demands, while employing commonly-available power electronics circuits.